FIG. 1 shows a ducted fan gas turbine engine 10 comprising, in axial flow series: an air intake 12, a propulsive fan 14 having a plurality of fan blades 16, an intermediate pressure compressor 18, a high-pressure compressor 20, a combustor 22, a high-pressure turbine 24, an intermediate pressure turbine 26, a low-pressure turbine 28 and a core exhaust nozzle 30. A nacelle 32 generally surrounds the engine 10 and defines the intake 12, a bypass duct 34 and a bypass exhaust nozzle 36. The engine has a principal axis of rotation 31.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct 34 and exits the bypass exhaust nozzle 36 to provide the majority of the propulsive thrust produced by the engine 10. The core flow enters in axial flow series the intermediate pressure compressor 18, high pressure compressor 20 and the combustor 22, where fuel is added to the compressed air and the mixture burnt. The hot combustion products expand through and drive the high, intermediate and low-pressure turbines 24, 26, 28 before being exhausted through the nozzle 30 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 24, 26, 28 respectively drive the high and intermediate pressure compressors 20, 18 and the fan 14 by interconnecting shafts 38, 40, 42.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In current engines, the high-pressure turbine gas temperatures are hotter than the melting point of the available superalloys meaning significant amounts of cooling air is required to prevent failure and premature aging. An alternative technology which helps alleviate the need for cooling air is ceramic matrix composites, or CMCs as they are commonly known.
Generally, CMC materials consist of ceramic fibres embedded within a ceramic body. There are different materials available for the fibres and body. Two of the more promising materials for gas turbine engines are silicon carbide fibres within a body of silicon carbide, so-called SiC/SiC, and aluminium oxide fibres within an aluminium oxide body, which is referred simply as an oxide CMC.
CMCs generally offer superior temperature and creep resistant properties for gas turbine engines and have a considerably lower density than their superalloy counterparts making them ideal for aeroengines. Further, because they have a higher temperature tolerance, CMC materials require less cooling which acts to increase specific fuel consumption further.
To help increase the mechanical robustness of CMC materials, it is advantageous to use continuous long fibres and avoid joins within a component. However, this is not immediately compatible with current architectures and assemblies of gas turbine engine components in which the aerofoils are provided within an annular gas path. This is particularly so for nozzle guide vanes which are conventionally formed in arcuate sections having a vane with radially inner and outer platforms. In a metallic cast component, the vane and platforms can be made as a homogenous structure in a casting process, with the absence of any joints between the vane and platform. However, this approach is not possible with CMCs due to the required lay-up of the fibres which prevents the acute transition from the vane to the platform. Creating the vanes and platforms separately is not considered viable due to difficulties associated with the subsequent joining of the separate components.
The present invention seeks to provide a nozzle guide vane assembly for a gas turbine engine and a method for creating the same.